Composite laminate structures are used extensively in the transportation industry for their lightweight and load bearing performance. As used herein, the transportation industry may include aerospace, marine, rail, and land-based vehicular applications. An aircraft, for example, may include stringers in an aircraft wing, stiffeners, and spar structures that are all formed of composite laminate structures.
A composite laminate structure may include a skin and a core. The core may be formed of end-grain balsa wood, a honeycomb of metallic foil or aramid paper, or of a wide variety of urethane, PVC, or phenolic foams, or the like. The skin may be constructed of multiple layers or laminates of a polymer matrix fiber composite, such as carbon fiber reinforced plastics (CFRP). Elements of the composite laminate structure may be bonded together, such as with an adhesive.
Typical failures in laminate structure can result from core failure under compressive forces or in shear or, more commonly, from a failure of the bond or adhesive capability between the core and the composite laminate skin, or between layers of the laminate skin. Other failures, depending on loading may include crimpling of one or both skins, bending failure of the laminate structure, or failure of the edge attachment means from which certain loads are transferred to the laminate structure.
Because of the failure modes noted above, reinforcements have been proposed that extend out-of-plane or normal to the planes of the skins. This is sometimes called the “Z” direction, as it is common to refer to the coordinates of the laminate skins as falling in a plane that includes the X and Y coordinates. Thus the X and Y coordinates are sometimes referred to as two-dimensional composite or 2-D composite. This is especially appropriate as the skins are often formed of fiber fabrics that are stitched or woven and stacked on top of each other to form plies or layers of a composite.
Conventional bonding methods are inadequate for certain types of laminates, may not significantly reinforce the laminates, and/or are overly burdensome to use during construction of the composite laminate structure. For example, Z-pinning has been proposed in which carbon pins are inserted at angles to fasten flanges of joined laminates. During this process, the ends of the pins protruding from the top surface of the laminate skin must be removed, which may weaken the pins due to micro-cracking. Additionally, the pins provide negligible, if any, reinforcement of the laminates. Another proposed approach is stitching the laminates together to bond them together and provide reinforcement. Stitching, however, cannot be performed on pre-impregnated laminates, as the impregnated resin will clog the sewing mechanism used to insert the stitch.
Therefore, it would be advantageous to have a method and apparatus that takes into account one or more of the issues discussed above.